NASA Technical Memorandum 100756

Development and Integration

of the Capillary Pumped Loop

GAS and Hitchhiker Flight

Experiments

D. Butler and R. McIntosh

April 1990

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NASA Technical Memorandum 100756

Development and Integration

of the Capillary Pumped Loop

GAS and Hitchhiker Flight

Experiments

D. Butler and R. McIntosh

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and

Space Administration

Goddard Space Right Center

Greenbelt, MD

1990

PREFACE

The Capillary Pumped Loop (CPL) is a thermal control system with high density heat acquisition and

transport capability. A small spaceflight version of the CPL was built and flown as a GAS experiment on

STS 51-D in April 1985 and STS 5143 in June 1985, and as a Hitchhiker-G experiment on STS 61-C in

January 1986. The purpose of the experiments was to demonstrate the capability of a capillary pumped

system under micro-gravity conditions for use in the thermal control of large scientific instruments, advanced orbiting spacecraft, and space station components. The development, integration, and test activities

of the CPL are described.

iii

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INTRODUCTION

The development of larger spacecraft and the space station has resulted in the need for new thermal

control systems that can handle large amounts of power and carry heat over long transport distances. Twophase systems arc being developed that require substantially less weight and power than existing singlephase systems. Two-phase thermal control systems can also maintain nearly constant temperatures over

wide power ranges, and can be used to supply heat as well as remove it.

One type of two-phase control system is the Capillary Pumped Loop (CPL). The CPL is unique in that

it has no moving parts. The original CPL concept was pioneered by F.J. Stenger of NASA/Lewis in the mid1960's (ref. 1) and has been under development at Goddard since the late 1970's. A schematic of the CPL

is shown in Figure 1. The evaporator contains a porous wick material (high density polyethylene, Porex)

which produces the pumping action in the closed loop system via capillary forces. The ammonia liquid

is drawn through the wick to the metallic shell of the evaporator where it vaporizes and then travels to

the condenser, thus transporting the heat via the latent heat of vaporization. The heat is removed at the

condenser and the vapor is returned back to a liquid state. The liquid then returns to the evaporator pumps

to continue the cycle.

Another important feature of the CPL is the two-phase reservoir. By controlling the reservoir temperature, the saturation temperature of the loop is controlled as well. This means that the evaporators stay at

a relatively constant temperature regardless of the heat load or condenser temperature variation. The evaporator temperature can be varied simply by raising or lowering the reservoir temperature to the desired

level. The reservoir also provides for automatic fluid inventory control in the loop by supplying or storing

ammonia liquid as required. The CPL also has an isolator located near the evaporator pump inlets. The isolator allows the evaporators to function individually within the loop and prevents vapor from flowing out

" of a non-operational evaporator into other evaporator pumps.

The CPL system has been undergoing development and testing for the past several years (see ref. 2).

Ground systems have demonstrated heat transport capabilities ranging from 100 watts to 25 kilowatts over

distances of 10 meters. Verification of the CPL under micro-gravity conditions was also pursued_ since

fluid behavior and evaporative and condensing heat transfer coefficients can be largely influenced by gravitational forces.

CPL/GAS FLIGHT EXPERIMENT

The next step in CPL development was the construction of a micro-gravity experiment. Drop tower

and low-g aircraft flights were considered but rejected because they only offered short durations at low

gravity (25 seconds or less). Since CPL startup requires 2 to 3 minutes, another approach was required.

Review of shuttle carrier options showed that the Get Away Special (GAS) carrier should be chosen for

the initial experiment due to its low cost, ease of integration, and frequent flight opportunities. The GAS

carrier consists of a cylindrical container with internal dimensions of 51-cm diameter by 71-cm long. GAS

experiments must be completely self-contained with no shuttle services available. Power. data. and control

functions must be carried along as part of the experiment. Nonetheless, the GAS carrier is well-suited for

small, basic research experiments at the proof-of-concept phase of their development.

The CPL/GAS experiment was developed utilizing existing hardware where possible (see Figure 2).

The support structure and battery were identical to those used for the STS-3 GAS Flight Verification Payload. The electronics and tape recorder were flown previously on the Atomic Oxygen Monitor GAS experiment flown on STS-8 and STS-11. Use of existing hardware can significantly reduce costs and shorten

development schedules.

Unfortunately, the space available for the CPL experiment was constrained due to the volume requirements of the battery and electronics. Nonetheless, a working mini-CPL was developed that mounts directly

to the GAS top plate between the structural support struts. It measured approximately 36 cm by 36 cm by

10-cm high and maintained most of the features of the large ground systems. Figure 3 shows the mini-CPL

experiment both before and after installation of the heaters, wiring, instrumentation, and electronics.

The CPL/GAS has two evaporator pumps mounted in parallel, with heaters attached directly to their

outer surfaces to provide the experiment heat load. The system was charged with approximately 150 grams

of ammonia. A single condenser tube, a temperature controlled two-phase reservoir, and an isolator are also
includedin this system.The boxesseenin Figure3 containadditionalelectronicsfor reservoirtemperaturecontrol andover-temperaturelimitstatsto preventoverheatingof theexperiment(viaheatercutoff).

The entire system was mounted to a 7-kg condenser plate which was attached directly to the I 1-kg GAS

container top plate to provide the system radiator/heat sink. The major portions of the experiment were

constructed of aluminum due to its light weight and high thermal conductivity. The reservoir and isolator

were fabricated from stainless steel due to its higher strength.

The CPL required 220 watts of power when it was in operation, with the power supplied by a Yardney

LR130 silver zinc battery. The battery was quite large, with a weight of 50 kg and external dimensions of

29 cm by 20 cm by 46 cm. It supplied 4 kilowatt-hours of energy to the CPL/GAS experiment.

The CPL/GAS experiment is shown fully assembled in Figure 4. The CPL was covered with a multilayered insulation blanket (MLI) which thermally isolates it from the battery and electronics. A thermostatically controlled heater was used on the battery to maintain its temperature above its lower limit of

0 ° C during potential cold case operations. The electronics box was covered with high-emittance Kapton

tape to radiatively dissipate its internally generated heat, which was approximately 12 watts. The GAS

container was flown without the insulating end cap used on most GAS payloads because of the large

heat rejection requirements of the CPL. The top plate served as the radiator for the experiment and its

outer surface was coated with silver Teflon to enhance its heat rejection capability.

Experiment sequencing was accomplished with the use of an electronic clock and pre-programmed,

hardwired memory. The memory was a commercially available bi-polar fusible link 8K ROM built by

the Harris Corporation. Experiment data was written onto a small Lockheed digital tape recorder conrained in the large electronics box. The data included 32 thermistor readings, power levels, calibration

voltages, and command status. Data was taken once a minute throughout the 5-day mission.

CPL/GAS TESTING AND FLIGHT QUALIFICATION

Since the CPL is a closed system containing ammonia, it is a pressure vessel and therefore subject to

special design requirements as specified by the NASA safety office. These include design to a 2428-psi pressure for all of the pressure system components. The .design pressure level is unique for each System, depending on the pressurant and the predicted maximum system pressure. The maximum pressure corresponds to

a worst-case expected maximum temperature of 80* C, which only occurs during a shuttle abort re-entry

and landing. The ammonia pressure at 80* C is 607 psi, and the CPL was designed to withstand 4 times that

pressure, or 2428 psi. Also, two identical CPL units were fabricated and then burst tested in order to prove

the design. Previously, there was more than one method that could be used to flight qualify pressure vessels

(see reference 3). We chose the least expensive of the options available. Now, however, a newer version of

the NASA safety document, NHB 1700.7B, has been prepared that restricts pressure vessel qualification to

specific portions of MIL-STD-1522A. This will significantly impact the design and test of future pressurized

systems for shuttle flight.

The CPL/GAS was subjected to testing which included a vibration test. thermal vacuum tests, and

extensive functional tests. A workmanship-level vibration test (at 6.4 g's RMS) was conducted to verify the

structural integrity of the experiment and to insure that it would endure the vibration loads that might be

encountered during the flight. Qualification testing of the structure was not required since the support

structure was previously qualified for earlier GAS flights. Detailed structural analysis was performed on all

of the new components that were built for CPL/GAS.

A thermal vacuum test was performed to insure proper operation of the CPL under extreme temperature conditions and the vacuum environment of space. The CPL/GAS container was situated upside-down

in the chamber to allow for proper operation of the CPL in the one-g earth environment. Temperature

variations were accomplished with a thermally controlled cold plate. It served as a direct radiative heat

sink for the GAS container top plate, which was the heat sink for the CPL experiment. The GAS container

also had to be levelled so that gravity effects on the CPL would be minimized.

The thermal vacuum test was conducted over a 6-day interval. The first 2 days of the test included

an experiment cooldown and cold case startup. The electronics were allowed to cool to -2 ° C and the

condenser cooled to -20 ° C for a cold start check. These levels corresponded to the predicted cold case

startup conditions. During the next 3 days of the test, a mission simulation was conducted. The thermal

environment (chamber and cold plate temperature) was set at -10 ° C. corresponding to the expected
shuttle payloadbay temperaturesfor the earthviewingcase.The missionprofile includedexperiment

heatercycleswith powerlevelsup to 220watts(110 watts perpump)for operatingtimesup to 1 hour,

followed by cooldownperiodslastingapproximately9 hours.Theseoperationtimeswerebasedon the

thermalanalysisof theCPL/GAS.

TheCPLsaturationtemperaturewascontrolledvia the reservoirtemperaturewhichwassetat 29° C.

This temperature was selected based on a number of factors. The temperature level of future on-orbit

thermal control systems will probably be near room temperature. Also, most of the ground tests on CPL

systems have been run near this temperature level. Finally, the ability to reject heat from the GAS container improves with higher radiator temperatures. Therefore, 29 ° C was selected as the best compromise

of all of these factors.

Prior to the start of each cycle, the CPL condenser was allowed to cool to approximately 5° C, then

the CPL evaporator heaters were activated. Since the heater power input exceeded the heat rejection capability of the GAS top plate, the condenser temperature increased. When the condenser temperature reached

the CPL operating temperature of 29* C, it could no longer absorb any more heat and the system was shut

down and again allowed to cool. The heater cycle was then repeated after the condenser cooled to about

5° C. A lower condenser temperature would have allowed for a longer experiment duration. However, the

cootdown portion of the cycle increases substantially as the shuttle environment (heat sink) temperature

of -10 ° C is approached. Therefore, 5° C was established as the best compromise between CPL operation

time and condenser cooldown duration. The mission simulation conducted as part of the test verified the

heatup and cooldown times predicted by the thermal analysis.

The last part of the thermal vacuum test was a hot case operational check, with the environment set

at 30 ° C. This test verified operation of the electronics at the expected hot case temperature. The CPL

heaters could not be activated because they require a heat sink at a temperature below the saturation temperature of 29 ° C.

A low-cost, functional test setup was also pursued so that the CPL could be further tested at length.

The CPL/GAS was again oriented upside-down, but now the GAS top plate rested on a continuously cooled

cold plate that removed the heat from the experiment via conduction. This provided more test time since

the top plate was cooled sufficiently to allow for continuous heater power application with no cooldown

time needed. Although this setup was not a realistic simulation of the Shuttle environment, it did allow for

low-cost, long-term experiment testing.

The functional testing continued for a total run time of approximately 8 weeks. Additional power profiles were developed and flight simulations were conducted. The value of testing cannot be overstated, especially for experiments dealing with new systems and technology development.

All shuttle payload organizations are required to submit detailed information about their experiments

in the form of safety review packages. Three to four safety reviews are held throughout the design and integration process to insure that the experiments are safe for flight on the shuttle and will not cause any

damage to the orbiter or crew. Every experiment component must be identified and design drawings,

parts and materials lists, and supporting analyses must be submitted. Quality Assurance requirements must

also be met, such as maintenance of certificates of material conformance and fabrication and test logbooks. The volume of documentation can often exceed the actual experiment volume by several orders of

magnitude.

The parts and materials selected for spaceflight use are subject to unique requirements. The materials

must be acceptable in terms of flammability, toxicity, freedom from stress corrosion cracking, vacuum

compatibility, and outgassing characteristics. Lists of acceptable materials are available to experiment

designers. All component parts must be vacuum compatible and able to withstand severe temperature

extremes and high vibration loads. Commercial parts seldom meet these requirements, so special military

standard or spaceflight quality parts are used instead. The prices for these items is often at least ten times

the cost of the corresponding commercial part.

One can readily understand the high cost of spaceflight when reviewing the myriad of special requirements that have to be met in order to have a successful experiment. The CPL/GAS cost was estimated to be

in excess of $500,000.00, and yet it was considered relatively "cheap" when compared to other flight

experiments. Although the costs are high, the rewards are also high when an experiment performs successfully and returns data on a new technology.
CPL/GAS FLIGHT

The first flight of the CPL/GAS experiment occurred in April 1985 on STS-51D. Unfortunately, the

GAS batteries that actuate the relay to energize the experiment failed, so the CPL could not be turned on

during the flight. The failure was due to a bad batch of batteries that failed under a combination of vacuum

and cold temperatures, even though they had passed qualification testing. The GAS project solved the problem by enclosing the batteries in a hermetically sealed box for all subsequent GAS flights. Fortunately for

the CPL/GAS project, a reflight opportunity only 2 months later was available on STS-51G in June 1985.

On this flight the GAS relays operated satisfactorily and the CPL experiment was very successful. The CPL

operated for the entire 120-hour mission and 13 power cycles worked as planned.

An example of a power profile from the flight is given in Figure 5. As heat was added to the evaporators their temperatures increased rapidly to the saturation temperature as controlled by the reservoir. At

that point, ammonia was vaporized and capillary pumping was started. The evaporator temperatures then

stabilized near the saturation temperature until the heaters were shut off at the end of the cycle. The evaporators then cooled slowly until the cycle was repeated later on. The CPL/GAS experiment results are presented in detail in ref. 4.

The CPL/GAS flight demonstrated that a capillary pumped system can perform well in micro-gravity.

No substantial differences between micro-g and one-g performance were noted. This was the first flight

of a thermal control system of this type and the first shuttle experiment from the Space Station Advanced

Development Program.

CPL/HH-G FLIGHT EXPERIMENT

A CPL follow-on mission on the Hitchhiker-Goddard (HH-G) was undertaken immediately after the

successful CPL/GAS flight. The Hitchhiker carrier has additional capabilities when compared to the GAS

carrier. These include shuttle power, real-time ground command and data, and additional volume and

weight.

The same mini-CPL hardware developed for the GAS flight was reflown on HH-G, with some modififications. The power per evaporator was increased from 110 watts to 400 watts (800 watts total) so that

the CPL could be tested at its maximum transport capability. The CPL was again flown in a GAS container, although no battery was needed. It was mounted to a specially designed 63-kg GAS container top plate

in order to accommodate the increased heat load. The insulation was removed from the sides of the contamer as well to increase the radiative heat rejection capability. Figure 6 shows the CPL/HH-G experiment

mounted along with the HH-G avionics and a PACS particle analysis experiment in the shuttle bay. Together they comprised the first HH-G carrier payload, with the flight aboard the Space Shuttle Columbia

on STS-61C in January 1986.

Testing of the CPL/HH-G included functional, vibration, thermal vacuum, and electro-magnetic interference (EMI) tests. The EMI tests were not required for GAS since the experiment was not electrically

connected to the shuttle. The tests are needed for HH-G since the shuttle provides electrical power, data.

and command capability. The EMI tests involve measurement of conducted and radiated electrical emissions from the experiment, and the experiment's susceptibility to emissions from the shuttle. The purpose

of the test is to insure that the experiment and shuttle do not interfere with one another electrically.

CPL/HH-G power profiles were defined during the functional tests. The profiles were similar to the

CPL/GAS profiles, but at a substantially higher power level. The power cycles again consisted of 30 to 60

minutes of heater power activation followed by a long cooldown time. An upper power limit of 600 watts

for the CPL-HHG was realized as a result of these tests. The thermal vacuum tests were used to verify

proper operation in vacuum under a simulated shuttle thermal environment. Hot and cold experiment operations were verified and a mission simulation was conducted.

The CPL/HH-G flight was very successful and again demonstrated CPL operation in micro-gravity.

Thirty-eight experiment power cycles were conducted over a 5-day mission. The CPL was monitored

around the clock and heater power levels were controlled by ground command. This level of flexibility

greatly enhanced the experiment data return. Additional power cycles were accomplished over the baseline

plan when the shuttle's stay on orbit increased by two days. New power profiles were developed in real

time based on the results obtained from previous cycles. Power profiles were adjusted in response to panic-
ular shuttlethermalattitudes.A sampleof theCPL/HH-Gdataisgivenin Figure7. TheevaporatortemperatureswerehigherthanontheGASexperimentdueto themuchhigherpowerlevelsusedonHH-G.Again.

ref.4 providesadetailedsummaryof theCPL/HH-Gflightresults.

TheCPL/HH-Gflight againdemonstratedsimilarCPLperformancein boththemicro-gandone-genvironments.With HH-G,upperandlowerpowerlimits wereestablishedfor the CPLflight unit. Postflight

testingduplicatedin-flight resultsandshowedthat theCPLhadnot degradedin performanceasaresultof

themission.

CONCLUSIONS

The CPL GAS and Hitchhiker flights successfully demonstrated the operation a capillary pumped twophase thermal control system in micro-gravity. A much larger CPL flight experiment is planned that will

demonstrate a heat transport capability of 1200 watts over a distance of 15 meters. This larger system is

a prototype of the CPL thermal control systems that will be used on future platforms and spacecraft.

The design, integration, and test of shuttle flight experiments is a somewhat arduous task due to all

of the special requirements of spaceflight. However, the excitement of flying an experiment in space and

the rewards of a successful mission make it all worthwhile.
REFERENCES

l° F.J. Stenger, "'Experimental Feasibility Study of Water-Filled Capillary-Pumped Heat-Transfer Loops",

NASA TM X-1310, NASA Lewis Research Center, Cleaveland, Ohio, 1966.

E.J.Kroliczek, J.Ku, S.Ollendorf, "Design, Development, and Test of a Capillary Pump Loop Heat

Pipe", AIAA-84-1720, AIAA 19th Thermophysics Conference, June 25-28, 1984.

.

"Safety Policy and Requirements, For Payloads Using the Space Transportation System (STS)",

NASA NHB 1700.7 Rev.A.

°

J. Ku, E.J. Kroliczek, D. Butler, R.B. Schweickart, R. Mclntosh, "Capillary Pumped Loop GAS

and Hitchhiker Flight Experiments", AIAA/ASME 4th Joint Therrnophysics Conference, June 2-4,

1986.
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Report Documentation Page Nalo_alA_nautcs ar¢l

:3oaceAclr_r_s_raion

1. Report No.

NASA TM-I00756

4. Title and Subtitle

2. Government Accession No. 3. Recipient's Catalog No.

5. Report Date

Development and Integration of the Capillary

Pumped Loop GAS and Hitchhiker Flight Experiments

7. Author(s)

D. Butler and R. Mclntosh

9. Performing Or_niza6onNameandAddress

Goddard Space Flight Center

Greenbelt, Maryland 20771

12. Sponsoring A_ncyNameandAddress

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

April 1990

6. Pedorming OrganizationCode

732

8. Pedormi_ Organization Report No.

90B00081

10. Work Unit No.

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstm_

The Capillary Pumped Loop (CPL) is a thermal control system with high density

heat acquisition and transport capability. A small spaceflight version of

the CPL was built and flown as a GAS experiment on STS 51-D in April 1985

and STS 51-G in June 1985, and as a Hitchhiker-G experiment on STS 61-C

in January 1986. The purpose of the experiments was to demonstrate the

capability of a capillary pumped system under micro_gravity conditions for

use in the thermal control of large scientific instruments, advanced

orbiting spacecraft, and space station components. The development, integration, and test activities of the CPL are described.

17K_ Wo_s(Suggesled_ Author(s))

Capillary Pumped Loop, G_t Away Special,

Hitchhiker-G, Thermal Control, MicroGravity, Two-Phase Heat Transfer,

Reservoirs, Space Shuttle, Space Station,

Polar Platforms, Evaporators

18. DistnbutionStatement

Unclassified - Unlimited

Subject Category 18

19. S_urityClassif.(ofth_report)

Unclassified

20. S_uriW Classif(ofthispege)

Unclassified

21. No. of pages ?'). Price

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Space Administration

Washington, 0 .C.

20546

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Space Adrninist ration

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